Turbine Blade Having Platform Cooling Holes

ABSTRACT

A turbine blade having a plurality of cooling holes which extend from an outside edge of the platform to a cooling passage formed within the turbine blade and a method of limiting the formation of cracks in the platform of the blade are provided. The plurality of cooling holes in the platform are formed at an approximate angle of 45° to the outside edge of the platform and are formed at the approximate mid-point of the thickness of the platform. The cooling holes are generally cylindrical in shape and have a diameter of approximately 50% of the platform thickness.

FIELD OF THE INVENTION

The present invention relates generally to techniques for reducing cracks in platforms of gas turbine blades and more specifically to a turbine blade having a plurality of cooling holes formed in the platform.

BACKGROUND

The turbine section of gas turbine engines typically comprises multiple sets or stages of stationary airfoils, known as nozzles or vanes, and moving airfoils, known as rotor blades or buckets. FIG. 1 illustrates a typical turbine rotor blade or turbine blade 100 found in the first stage of the turbine section, which is the section immediately adjacent to the combustion section of the gas turbine and thus is in the region of the turbine section that is exposed to the highest temperatures. Known problems with such blades 100 are the formation of multiple cracks and, when present, the delamination of a thermal barrier coating (TBC) in the platform region 104 due to the heat stresses in this region of the blade. In some cases the cracking is so severe that it results in breakage and separation of a substantial portion of the platform 104 on the pressure side of the blade 100, leading to the early retirement of the blade. In order to prevent this early retirement and to extend blade operational lifetime, various approaches have been proposed.

In one such solution, an undercut is machined into the blade platform along the pressure side of the blade. This proposed solution purports to reduce the total stress level in this region of high stress. This approach has been implemented on both turbine and compressor blades as both a field repair and a design modification. If a stress reduction is achieved in the platform region, the concern is whether the undercut results in a high stress within the grooved region where material is removed. In other words, the success of the strategy depends on whether a stress reduction in an existing high-stress region can be achieved without creating a new area of high stress within the blade.

There are two primary concerns raised with platform undercuts. First, whether the undercut will be effective in reducing the stress in the platform. Second, whether the stress concentration occurring in the undercut will be so high that it offsets the benefit of the undercut to the platform region. Undercut solutions have had difficulty striking a balance between these two concerns. It is desired to have a solution which effectively reduces the stress in the platform and, thereby, the potential for formation of cracks, TBC delamination, and, in the worst case, breakage and separation of significant portions of the platform altogether, and which does not add additional stresses to the blade. The present invention seeks to solve this problem.

SUMMARY

In one embodiment, the present invention is directed to a turbine blade and limits platform cracking. The turbine blade has of an airfoil connected to a platform in the blade root region. The airfoil and the platform share a common cooling passage, which may include one or more cooling channels or paths. The turbine blade configuration limits platform damage, including but not limited to cracking, removal of the TBC layer, and breakage and loss of blade material, through the influence of at least one cooling hole which extends from an outside edge of the platform, through the platform and to the common cooling passage. In one embodiment, multiple cooling holes are formed in the platform at an angle to the outside edge of the platform which is approximately 90° and in another embodiment at approximately 45°. In one embodiment, the common cooling passage has one or more serpentine cooling circuits, and each of the cooling holes extends to a distinct channel within the serpentine cooling circuit. In another embodiment, the common cooling passage includes a plurality of generally parallel cooling veins extending through the platform and airfoil to the airfoil tip, which may be formed by a Shaped-Tube Electolytic Machining (STEM) drilling process for example. In this case, each of the platform cooling holes extends through the platform from the platform edge to a distinct parallel cooling vein in the blade. In one embodiment, the platform cooling holes are formed at the approximate midpoint of the defined thickness of the platform. The cooling holes may be generally cylindrical in shape with a diameter approximately 50% of the platform thickness.

In another embodiment, the present invention is directed to a method of limiting damage to the platform of a turbine blade having an airfoil connected to the platform in a blade root region. The method includes the step of forming at least one cooling hole in the platform which extends from an outside edge of the platform to a cooling passage within the platform which passes through and is shared with the airfoil. In one embodiment, multiple cooling holes are formed in the platform. In a more specific embodiment, there may be four cooling holes formed in the platform. The cooling holes can be formed at the angles, having the location and geometric dimensions as described in the immediately preceding paragraph. The cooling holes can be formed by a number of known processes, but in at least one embodiment is formed by an electro-discharge machining (EDM) process.

The present invention has application in both the manufacture of a new turbine blade as well as in the repair of an existing turbine blade not having cooling holes formed in the platform region thereof. In the latter case, the one or more cooling holes can be formed without having to remove any existing TBC layer that may have been applied on the blade.

BRIEF DESCRIPTION OF THE DRAWINGS

The following drawings form part of the present specification and are included to further demonstrate certain aspects of the present invention. The present invention may be better understood by reference to one or more of these drawings in combination with the description of embodiments presented herein. However, the present invention is not intended to be limited by the drawings.

FIG. 1 is a perspective view of a prior art turbine blade having a portion of its platform eroded.

FIG. 2 is a perspective view of a turbine blade in accordance with the present invention illustrating a plurality of cooling holes formed in its platform.

FIG. 3 is a cross-sectional view of the platform of the turbine blade taken across line 3-3 shown in FIG. 2 illustrating the platform cooling holes of the present invention communicating with the corresponding distinct cooling pathways of a serpentine cooling circuit.

FIG. 4 is a cross-sectional view of the platform of the turbine blade taken across the same line 3-3 shown in FIG. 2 illustrating the platform cooling holes of the present invention communicating with a corresponding plurality of generally parallel cooling veins formed in the airfoil and platform.

FIG. 5 is a cross-sectional view of a turbine blade with two separate serpentine cooling passages in the airfoil, according to various embodiments of the present invention.

FIGS. 6A and 6B are plots of the distribution of heat transfer coefficient (HTC or film coefficient) along the leading and trailing serpentine cooling circuits (shown in FIG. 5), respectively, as a function of the distance from the air inlets at the base of the blades, to each corresponding (leading or trailing) cooling circuit, according to various embodiments of the present invention.

FIGS. 7A through 7D are plots of the film coefficient and cooling air temperature along the length of the platform cooling holes as shown in FIG. 3, according to various embodiments of the present invention. The film coefficients and temperatures are shown as a function of distance from the point where the platform cooling holes join the serpentine cooling circuits for each of the four platform cooling holes.

FIGS. 8A and 8B are perspective views of the turbine blade with (FIG. 8A) and without (FIG. 8B) the platform cooling holes according to various embodiments of the present invention, respectively, illustrating the metal temperature distribution on the surface of the entire blade.

FIGS. 9A and 9B are perspective views of the turbine blade with (FIG. 9A) and without (FIG. 9B) the platform cooling holes according to various embodiments of the present invention, respectively, illustrating the temperature distributions in the region of these blades proximate the platform cooling holes. The left hand illustration shows the blade platform in perspective view from above and the right hand illustration shows the platform temperatures looking from below in each of the two figures.

FIGS. 10A and 10B are perspective views of the turbine blade with (FIG. 10A) and without (FIG. 10B) the platform cooling holes according to various embodiments of the present invention, respectively, illustrating the temperature distribution in the region of the blade proximate the juncture of the platform and trailing edge lowermost cooling hole.

FIGS. 11A and 11B are perspective views of the turbine blade with (FIG. 11A) and without (FIG. 11B) the platform cooling holes according to various embodiments of the present invention, respectively, illustrating the equivalent stress distributions in the platform region. The left hand illustration shows the sectioned blade and platform looking down from above while the right hand illustration shows the sectioned blade shank and platform looking up from below, in each figure respectively.

FIGS. 12A and 12B are perspective views of the turbine blade with (FIG. 12A) and without (FIG. 12B) the platform cooling holes according to various embodiments of the present invention, respectively, illustrating the axial stress distributions in the platform region. The left hand illustration shows the sectioned blade and platform looking down from above while the right hand illustration shows the sectioned blade shank and platform looking up from below, in each figure respectively.

FIGS. 13A and 13B are perspective views of the turbine blade with (FIG. 13A) and without (FIG. 13B) the platform cooling holes according to various embodiments of the present invention, respectively, illustrating the stress distributions proximate the juncture of the platform and lowermost, trailing edge cooling hole.

DETAILED DESCRIPTION

The present invention will now be described with reference to the following exemplary embodiments. Referring now to FIG. 2, a turbine blade in accordance with the present invention is shown generally by reference number 200. The turbine blade 200 has three primary sections, a shank 202, which is designed to slide into a disc on the shaft of the rotor (not shown); a platform 204 connected to the shank 202; and an airfoil 206 connected to the platform 204. Platform 204 connects to shank 202 at a lower surface of the platform 204, and to airfoil 206 at an upper surface of the platform. Platform 204 has a thickness defined by the distance between the lower surface and the upper surface. Moreover, platform 204 has four outside edges, which are generally orthogonal to the lower surface and the upper surface. Generally, during the blade 200's initial manufacture, the shank 202, platform 204 and airfoil 206 are all cast as a single part.

The airfoil 206 is defined by a concave side wall 210, a convex side wall 208, a leading edge 212 and an opposite trailing edge 214; the leading and trailing edges being the two areas where the concave side wall and convex side wall meet. The airfoil 206 has a root 216 which is proximate the platform 204 and a tip (or shroud) 218 which is distal from the platform. As with prior art turbine blades, air is supplied to the inside cavity (not shown) of the airfoil 206 from the compressor to cool the airfoil. The cooling air may exit through a plurality of cooling holes 220, at least some of which may be formed in the trailing edge 214.

In accordance with the present invention, the platform 204 has a plurality of cooling holes 230 formed therein on the pressure (concave) side of the airfoil 206, which is the region of the platform that is susceptible to high stresses and often cracks, including delaminating of coating, when present, and separation or breakage of blade base material in extreme cases. In one embodiment, four such cooling holes 230 are formed in the platform 204. The platform cooling holes 230 may be formed by an EDM process. Alternatively, the platform cooling holes 230 can be formed via STEM process or electro-chemical (ECM) drilling process or other similar machining process. The process utilized to form the cooling holes 230 may be selected to avoid removal of the TBC layer formed on the turbine blade. In one embodiment, the platform cooling holes 230 are generally cylindrical in shape, with center axes generally parallel to the lower surface and the upper surface. The cross-section of a platform cooling hole 230 at an outside edge of the platform 204 may span approximately 50% of the platform thickness, or the platform cooling holes 230 may have a diameter of approximately 50% of the thickness of the platform 204. The platform cooling holes 230 may also be formed at the approximate mid-point of the thickness of the platform 204, i.e., the centers of the cross-section of the platform cooling holes 230 at the outside edge of the platform 204 are aligned at the mid-point of the platform thickness so that an equal amount of platform material is left above and below the platform cooling holes 230.

In one embodiment, the platform cooling holes 230 are formed at an angle to the outside edge of the platform 204 into which they are formed, which is best seen in FIG. 3. The platform cooling holes 230 intersect the cooling cavity or passage 240, which platform 204 shares with the airfoil 206 and which is fed by cooling air from the compressor section of the turbine (not shown). In the embodiment shown in FIG. 3, the common cooling passage 240 is defined by a pair of serpentine cooling circuits, namely a leading serpentine cooling circuit 242 and a trailing serpentine cooling circuit 244. In turn, each of the serpentine cooling circuits is defined by a plurality of generally parallel channels or pathways 246. The orientation and location of the serpentine cooling circuits are shown in cross section in FIG. 5. As illustrated in FIG. 3, the platform cooling holes 230 form an angle a with the edge of the platform 204 which is approximately 45°. Each of the platform cooling holes 230 is illustrated as extending to, and communicating with, a distinct cooling pathway 246. The cooling air thus flows from the compressor to the turbine blade 200 first through a cavity in the shank 202 (not shown), then through the cooling pathways 246 of the serpentine cooling circuits 242, 244 and then through the platform cooling holes 230 before exiting the turbine blade. As the cooling air flows through the platform cooling holes 230 it cools the platform 204, thereby preventing delamination of the TBC layer, formation of cracks, and, worse, breakage and separation of the platform in that region altogether.

In another embodiment (shown in FIG. 4), the platform cooling holes 230 are formed in the platform 204 at an angle and orientation within the thickness of the platform, but instead extend to, and communicate with, a corresponding plurality of generally parallel cooling veins 250 formed in the platform. In other words, the common cooling passage 240 in FIG. 4 is a plurality of discrete generally parallel cooling veins 250. The cooling veins 250 may be formed by a number of processes, but usually are formed by a STEM drilling process. The cooling veins 250 intersect a cavity (not shown) in the shank 202 of the turbine blade 200, which is fed by cooling air from the compressor (also not shown).

Without limiting the invention to a particular theory or mechanism of action, it is nevertheless currently believed that the overall cooling flow may increase and the internal cooling flow may be re-distributed as a consequence of adding the platform cooling holes 230. Table I lists the cooling mass flow which may occur as a result of adding the platform cooling holes 230 to an example first stage turbine blade with serpentine cooling passages.

TABLE I COMPARISON OF COOLING FLOW RATE Prior Art Blade with Platform Blade Cooling Holes Difference Leading Serpentine (lb_(m)/hr) 453 456 +0.7% Trailing Serpentine (lb_(m)/hr) 512 518 +1.2% Total (lb_(m)/hr) 965 974 +0.9%

As shown in the table, the cooling flow in the leading serpentine cooling circuit may be ˜0.7% more than the prior art blade configuration, and the cooling flow in the trailing serpentine cooling circuit may increase by ˜1.2%. The total cooling flow may increase only marginally (by ˜0.9%) with the drilling of four platform cooling holes 230. The cooling flow of the leading three platform cooling holes 230 may be 6.1, 5.8, and 6.5 pound mass per hour (lb_(m)/hr), respectively. For the 4th platform-cooling hole 230, which branches from the trailing serpentine passage, the flow rate may be 6.1 lb_(m)/hr. The total platform cooling flow may be 24.4 lb_(m)/hr, or about 2.5% of total cooling flow available to the bucket.

FIG. 5 shows separate serpentine cooling passages in the airfoil. The leading edge serpentine cooling circuit 252 cools the leading, front half of the blade and receives its cooling air from inlets 1 and 2, which are located at the base of the blade and lead into cavity 256. The trailing edge serpentine cooling system 254 cools the trailing, back half of the blade and receives its cooling air from inlets 3 and 4, which are located at the base of the blade and lead into cavity 258.

FIGS. 6A and 6B show the distribution of heat transfer coefficient (film coefficient) along the leading and trailing serpentine circuits, respectively, according to one embodiment of the invention. It can be seen that drilling four platform-cooling holes 230 may have a minimal impact on the original cooling of the main internal flow. The computed cooling flow parameters for each platform cooling hole are shown in FIGS. 7A-7D, respectively.

The resulting surface temperature distributions of the modified blade with platform cooling holes 230, according to one embodiment of the invention, and a prior art blade are shown in FIGS. 8A and 8B, respectively. As indicated by the results of the cooling flow analysis, the thermal response in the airfoil above the platform is basically unchanged when compared to the temperature distribution of the original design configuration. As a consequence of the insertion of four parallel platform cooling holes 230, a substantial reduction of temperature was predicted in the region encompassing the platform cooling holes 230. The peak temperature predicted on the pressure side of the platform was significantly reduced from approximately 1800° F. for the original design to 1600° F. for the modified platform, e.g. a drop of about 200° F. This is illustrated in FIGS. 9A and 9B. As indicated by this drop, the platform cooling holes 230 are effective as they extract fresh coolant air from the serpentine cooling circuit and provide maximum coverage possible over the pressure side region of the platform.

Further examining these results indicates that the benefit of the proposed platform cooling strategy is likely to be at least twofold. Through the additional convective cooling and conduction, the gross reduction of the temperature in the platform region should favorably lower the temperature gradients near the juncture of platform and trailing edge lowermost cooling hole, which is particularly susceptible to cracking, as indicated in FIGS. 10A and 10B. Temperatures near the trailing edge lowermost cooling hole may be lowered by approximately 10° F.

Equivalent and axial stress distributions of the blade modified with platform cooling, according to one embodiment of the invention, are plotted in FIGS. 11A and 11B and FIGS. 12A and 12B, respectively. In the prior art turbine blades, there are large compressive stresses induced by platform curling due to the temperature gradients across the platform 204 and airfoil/shank region (under steady load). The excessive compressive stress at base load indicates a potential for substantial damage resulting from the out-of-phase thermal-mechanical fatigue (TMF) that would occur from each start-stop cycle. As shown in FIGS. 11A and 11B and FIGS. 12A and 12B, overall stress levels on the pressure side of the platform 204 may be reduced considerably by about 10-30%. At the free edge, near the exit of the platform cooling holes 230, the critical minimum principal stress may be reduced from 93 kilo-poind per square inch (ksi) to 62 ksi as a result of platform cooling modification. In the mid-span, the critical minimum principal stress may decrease from 111 ksi to 100 ksi. This relatively mild stress is localized and attributed to the thermal gradient across the platform-cooling hole 230. Nevertheless, lowering the metal temperature by 150° F.˜200° F. may significantly enhance the associated fatigue properties and, hence, increase the corresponding TMF life. TMF life may improve by as much as 200% by taking into consideration the fatigue property benefits resulting from the calculated temperature improvement (Table II). In addition, with a much lower stress predicted at the free edge near trailing edge of the blade, the fatigue crack propagation life may improve substantially comparing to the original design.

TABLE II COMPARISON OF STRESS RESULTS IN THE PLATFORM Critical Min. Principal % Change Estimated % Change of Stress (ksi) of Stress TMF Life Prior Art Blade 111 0%  0% Blade with Platform 100 −10% +200%* Cooling Holes *taking into account the temperature effect on TMF property

FIGS. 13A and 13B show the stress distribution in the lowermost cooling hole region after the platform cooling modification, according to one embodiment of the invention. As illustrated in the plot, the lower thermal gradient near the junction of airfoil trailing edge and platform favorably reduces the stress at the critical location from 83 ksi to 76 ksi, or a drop of about 8% (Table III). The corresponding TMF life may increase by ˜100% as a consequence of the platform cooling modification.

TABLE III COMPARISON OF STRESS RESULTS IN THE LOWERMOST COOLING HOLE Critical Max. Principal % Change Estimated % Change of Stress (ksi) of Stress TMF Life Prior Art Blade 83 0% 0% Blade with Platform 76 −8% +100% Cooling Holes

In summary, the platform cooling hole modifications of the present invention may be effective in both reducing the temperatures and stresses in the cooled platform region. Moreover, they may provide additional benefits in lowering the thermal gradient near the juncture of platform and trailing edge, and consequentially reduce the stress at the trailing edge lowermost cooling hole. Based on a comparison to the results of the baseline analysis, it is therefore considered as a viable design modification—to be utilized in the course of forming a new turbine blade—and/or recommended to implement during repair and refurbishment of blades.

The terms “holes,” “passages,” “veins,” “channels,” and the like are each used to describe conduits for the flow of air or other cooling fluid. The use of different words for the various conduits is not intended to be limiting in any way, but instead is to assist the reader in fully understanding the interrelation between the various conduits.

Therefore, the present invention is well adapted to attain the ends and advantages mentioned as well as those that are inherent therein. The particular embodiments disclosed above are illustrative only, as the present invention may be modified and practiced in different but equivalent manners apparent to those skilled in the art, having the benefit of the teachings herein. For example, as those of ordinary skill in the art will appreciate a different number of platform cooling holes 230 may be implemented, such platform cooling holes 230 may be formed at a different angle than that disclosed herein, and such platform cooling holes 230 may be oriented at a different location within the thickness of the platform. Furthermore, no limitations are intended to the details of construction or design herein shown, other than as described in the claims below. It is therefore evident that the particular illustrative embodiments disclosed above may be altered or modified and all such variations are considered within the scope and spirit of the present invention. Also, the terms in the claims have their plain, ordinary meaning unless otherwise explicitly and clearly defined by the patentee. 

1. A turbine blade having an airfoil connected to a platform in a root region, the airfoil and platform having a common cooling passage formed therein, the platform comprising at least one cooling hole which extends from an outside edge of the platform to the common cooling passage.
 2. The turbine blade according to claim 1, wherein a plurality of cooling holes are formed in the platform which extend from the outside edge of the platform to the common cooling passage.
 3. The turbine blade according to claim 2, wherein four cooling holes are formed in the platform which extend from the outside edge of the platform to the common cooling passage.
 4. The turbine blade according to claim 3, wherein the plurality of cooling holes are formed at an angle to the outside edge of the platform which is less than 90°.
 5. The turbine blade according to claim 4, wherein the angle at which the plurality of cooling holes are formed to the outside edge of the platform is approximately 45°.
 6. The turbine blade according to claim 5, wherein the common cooling passage includes a serpentine cooling circuit.
 7. The turbine blade according to claim 6, wherein each of the cooling holes extends to a distinct pathway within the serpentine cooling circuit.
 8. The turbine blade according to claim 5, wherein the common cooling passage includes a plurality of generally parallel cooling veins extending through the platform and airfoil.
 9. The turbine blade according to claim 8, wherein each of the cooling holes extends to a distinct parallel cooling vein.
 10. The turbine blade according to claim 1, wherein the platform has a defined thickness and the at least one cooling hole is formed at the approximate mid-point of the thickness.
 11. The turbine blade according to claim 1, wherein the at least one cooling hole is generally cylindrical in shape.
 12. The turbine blade according to claim 1, wherein the platform has a defined thickness, and the at least one cooling hole has a diameter of approximately 50% of the platform thickness.
 13. A method of limiting damage to a platform of a turbine blade having an airfoil connected to the platform in a root region of the airfoil blade, the airfoil and platform having a common cooling passage formed therein, the method comprising the step of forming at least one cooling hole in the platform which extends from an outside edge of the platform to the common cooling passage.
 14. The method according to claim 13, wherein a plurality of cooling holes are formed in the platform which extend from the outside edge of the platform to the common cooling passage.
 15. The method according to claim 14, wherein four cooling holes are formed in the platform which extend from the outside edge of the platform to the common cooling passage.
 16. The method according to claim 15, wherein the plurality of cooling holes are formed at an angle to the outside edge of the platform which is less than 90°.
 17. The method according to claim 16, wherein the angle at which the plurality of cooling holes are formed to the outside edge of the platform is approximately 45°.
 18. The method according to claim 17, wherein the cooling passage includes a serpentine cooling circuit.
 19. The method according to claim 18, wherein each of the cooling holes extends to a distinct pathway within the serpentine cooling circuit.
 20. The method according to claim 17, wherein the cooling passage includes a plurality of generally parallel cooling veins extending through the platform and airfoil.
 21. The method according to claim 18, wherein each of the cooling holes extends to a distinct parallel cooling vein.
 22. The method according to claim 13, wherein the platform has a defined thickness and the at least one cooling hole is formed at the approximate mid-point of the thickness.
 23. The method according to claim 13, wherein the platform has a defined thickness and the at least one cooling hole is generally cylindrical in shape and has a diameter of approximately 50% of the platform thickness.
 24. The method according to claim 13, wherein the at least one cooling hole is formed by an EDM process.
 25. The method according to claim 13, wherein the step of forming at least one cooling hole is performed in the course of forming a new turbine blade.
 26. The method according to claim 13, wherein the step of forming at least one cooling hole is performed in the course of repairing a turbine blade that has previously been in service.
 27. The method according to claim 25, wherein the step of forming at least one cooling hole is performed without removing any TBC layer formed on the turbine blade. 